Infrared suppressing exhaust system
A method of managing aircraft exhaust includes providing hot air at a hot air mass flow rate, providing cold air at a cold air mass flow rate, and mixing the hot air and the cold air at a variable hot air mass flow rate to cold air mass flow rate ratio, wherein the variable hot air mass flow rate to cold air mass flow rate ratio is selectively maintained independent of at least one of (1) a variation in the hot air mass flow rate and (2) a variation in a translational speed of the aircraft.
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STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH OR DEVELOPMENTNot applicable.
REFERENCE TO A MICROFICHE APPENDIXNot applicable.
BACKGROUNDExhaust systems, such as those associated with aircraft, may expel exhaust at undesirably high temperatures and/or may provide line of sight viewing of components that are heated to an undesirably high temperature. In some cases, the excessive temperatures may contribute to an undesirable heat signature perceptible by heat based sensing devices. In some cases, undesirably high temperature exhaust may contribute to undesired heating and/or scorching of a landing surface.
SUMMARYIn some embodiments of the disclosure, a method of managing aircraft exhaust is provided that comprises providing hot air at a hot air mass flow rate, providing cold air at a cold air mass flow rate, and mixing the hot air and the cold air at a variable hot air mass flow rate to cold air mass flow rate ratio, wherein the variable hot air mass flow rate to cold air mass flow rate ratio is selectively maintained independent of at least one of (1) a variation in the hot air mass flow rate and (2) a variation in a translational speed of the aircraft.
In other embodiments of the disclosure, an exhaust system is disclosed as comprising a mixer duct comprising a mixer duct internal space at least partially bounded by a mixer duct wall, a distributor at least partially disposed within the mixer duct internal space, the distributor comprising a distributor internal space at least partially bounded by a distributor wall and at least one distributor perforation through the distributor wall, the distributor perforation being configured to join the distributor internal space with the space between the mixer duct wall and the distributor wall in fluid communication, and an exhaust system controller configured to selectively affect a first air mass flow rate through the space between the mixer duct wall and the distributor wall.
In yet other embodiments of the disclosure, an exhaust system is disclosed as comprising a hot air source configured to selectively provide the exhaust system with air comprising a hot air temperature at a hot air mass flow rate and an exhaust system controller configured to selectively provide the exhaust system with air having a cold air temperature and a variable cold air mass flow rate, wherein the cold air temperature is less than the hot air temperature, and wherein the cold air controller is configured to selectively achieve a selected cold air mass flow rate independent of the hot air mass flow rate.
For a more complete understanding of the present disclosure and the advantages thereof, reference is now made to the following brief description, taken in connection with the accompanying drawings and detailed description:
It should be understood at the outset that although an illustrative implementation of one or more embodiments are provided below, the disclosed systems and/or methods may be implemented using any number of techniques, whether currently known or in existence. The disclosure should in no way be limited to the illustrative implementations, drawings, and techniques illustrated below, including the exemplary designs and implementations illustrated and described herein, but may be modified within the scope of the appended claims along with their full scope of equivalents.
In some cases, it may be desirable to provide an exhaust system that provides a variable ratio of cold air mass flow rate to hot air mass flow rate that is selectively variable selectively independent of engine speed, vehicle speed, and/or direction of expelled exhaust. In some embodiments of the disclosure, an exhaust system is provided that supplies an interior space of a distributor with an independently variable mass flow rate of cold air. In some embodiments of the disclosure, an exhaust system is provided that additionally supplies an independently variable mass flow rate of cold air into a space associated with an exterior of the above-mentioned distributor. In some embodiments of the disclosure, the mass flow rate of cold air supplied to the interior space of the distributor and/or the mass flow rate of cold air supplied to the space associated with the exterior of the distributor may be varied to mix with hot air supplied to the exhaust system to achieve a desired expelled exhaust temperature and/or a desired temperature of a component of the exhaust system. Such a reduction in temperature may substantially reduce and/or eliminate an infrared detectability of an aircraft 100 employing the presently disclosed exhaust system.
Referring to
Referring to
The nacelle 101 further comprises a longitudinal axis 114, a lateral axis 116, and a vertical axis 118. The longitudinal axis 114 generally extends longitudinally in a front-rear direction relative to the nacelle 101. The lateral axis 116 generally extends laterally in a left-right direction relative to the nacelle 101. The vertical axis 118 generally extends vertically in a top-bottom direction relative to the nacelle 101. The longitudinal axis 114, lateral axis 116, and vertical axis 118 intersect each other at an origin 120 and may generally be described as defining a three dimensional Cartesian coordinate system. The nacelle 101 further comprises a lateral bisection plane 122, a longitudinal bisection plane 124, and a vertical bisection plane 126. The lateral bisection plane 122 is generally coincident with the lateral axis 116 and the vertical axis 118. The longitudinal bisection plane 124 is generally coincident with the longitudinal axis 114 and the vertical axis 118. The vertical bisection plane 126 is generally coincident with the lateral axis 116 and longitudinal axis 114.
While the axes 114, 116, and 118 and bisection planes 122, 124, and 126 are generally defined to facilitate discussion of the nacelle 101, the location of the origin 120 relative to the nacelle 101 and the orientation of the axes 114, 116, and 118 relative to the nacelle 101 and/or to a primary direction of forward movement of the aircraft 100 may be described differently without impact to the functionality of the aircraft 100, the nacelle 101, and/or the components of the nacelle 101 disclosed herein. In other words, unless otherwise noted herein, the defined orientations of the axes 114, 116, and 118 and bisection planes 122, 124, and 126 are provided as a frame of reference against which the nacelle 101 and the components of the nacelle 101 may be consistently described.
Still referring now to
Still referring to
In some embodiments, the exhaust system 200 comprises an exhaust system controller 215 (see
Referring now to
Referring now to
Referring now to
The mixer duct 204 generally comprises a mixer duct interior space 242 substantially bounded about the longitudinal axis 114 by a mixer duct wall 244 that generally longitudinally terminates at a mixer duct inlet 246 and a mixer duct outlet 248. The mixer duct 204 is generally joined to the surface supply duct 210 that extends laterally rightward and longitudinally forward of the mixer duct wall 244 and comprises a surface supply duct interior space 250 in fluid communication with the mixer duct interior space 242. The mixer duct 204 further comprises mixer duct apertures 252 along a top portion of the mixer duct wall 244 for receiving cold air from the cavity supply duct 208. The cavity supply duct outlet 236 is generally shaped complementary to the mixer duct wall 244.
Referring now to
Referring now to
The hot air may be received into the exhaust system 200 and into the mixer duct interior space 242 within the tubular blade assembly 240 of the surface fan 216. In cases where the aircraft 100 is stationary and neither the cavity fan 212 nor the surface fan 216 are operated, the hot air may generally pass through the space between the distributor wall 254 and the mixer duct wall 244 and flow longitudinally rearward until the hot air exits the exhaust system 200 through the mixer duct outlet 248. In some embodiments, the cavity fan 212 and the surface fan 216 may be operated to introduce cold air into the distributor interior space 256 and the space between the tubular blade assembly 240 and the mixer duct wall 244, respectively. In some cases, the cavity fan controller 218 may control the cavity fan 212 to supply cold air to the distributor interior space 256 at a pressure sufficient to force cold air through at least one of the distributor perforations 260. By flowing cold air from the distributor perforations 260 while hot air and/or cold air passes longitudinally over the exterior of the distributor wall 254, film cooling may be achieved. Additionally, the surface fan controller 218 may be operated to control the surface fan 216 to induce cold air through the surface supply duct 210 and become entrained in the longitudinal air flow prior to entering the space between the distributor wall 254 and the mixer duct wall 244.
In some embodiments, utilizing one or more of the above-described cavity fan 212 and surface fan 216 to selectively supply cold air for mixing with hot air may provide an ability to select, achieve, and/or maintain a selected cold air mass flow rate to hot air mass flow rate ratio regardless of the translational speed of the aircraft, regardless of the orientation of the exhaust system 200 relative to a landing surface, and/or regardless of an engine 132 speed. While ducts and fans are disclosed above as providing the ability to selectively vary the cold air mass flow rate, in alternative embodiments, one or more of the exhaust system 200 components may be configured to provide a variable orifice and/or cross sectional area to alter pressures and/or mass flow rates through the variable orifice and/or cross-sectional area. Such variations in orifices, cross-sectional areas, and/or the cold air mass flow rate may affect an aerodynamic resistance to movement of the aircraft 100. For example, in some cases, selectively reducing an amount of cold air routed through the exhaust system 200, reducing an orifice size, reducing a cross-sectional flow area, and/or reducing a size of an external surface area of the exhaust system may selectively reduce an amount of power required to move the aircraft 100 during flight, such as flight while the aircraft 100 is in the airplane mode shown in
The above-disclosed exhaust system 200 is configured to specifically prevent visual observance of high temperature exhaust system 200 components, thereby reducing an amplitude of an infrared heat signature of the aircraft 100. Further, one or more of the above-described air mass flow rates may be selectively controlled in response to a sensed threat condition so that relatively more cold air may be selectively be mixed with the hot air to selectively enter a stealth mode or to further reduce an infrared heat signature. Still further, the ratio of cold air mixed with hot air may be controlled to reduce a temperature of air expelled from the exhaust system 200 so that the exhaust system 200 does not scorch, burn, or warp a landing and/or take-off surface. Additionally, in this embodiment, while the surface fan 216 comprises an electric outer ring or annular motor 238, any other type of drive may be utilized to power the surface fan 216, such as, but not limited to, a belt drive, a chain drive, an engine shaft, a gear drive, a hydraulic drive, and/or any other suitable drive system. The annular motor 238 may be desirable at least insofar as it allow location of sensitive bearings, electrical windings, and other heat sensitive surface fan 216 components radially outside the hot air flow path, thereby offering increased service life. Also, the generally angularly symmetrical shape of a plurality of the exhaust system 200 components may provide increased resistance to stress and/or strain failures due to high heat and vibratory conditions as compared to other generally box-shaped and/or rectangular components of some infrared signature suppression system. While the exhaust system 200 is described herein as applied to aircraft 100, in other embodiments, the exhaust system 200 may similarly applied to any other system comprising and engine and/or a source of hot air, such as, but not limited to, an airplane, a helicopter, a land vehicle, a water vehicle, a generator, and/or any other suitable system that may comprise a source of hot air.
Referring now to
At least one embodiment is disclosed and variations, combinations, and/or modifications of the embodiment(s) and/or features of the embodiment(s) made by a person having ordinary skill in the art are within the scope of the disclosure. Alternative embodiments that result from combining, integrating, and/or omitting features of the embodiment(s) are also within the scope of the disclosure. Where numerical ranges or limitations are expressly stated, such express ranges or limitations should be understood to include iterative ranges or limitations of like magnitude falling within the expressly stated ranges or limitations (e.g., from about 1 to about 10 includes, 2, 3, 4, etc.; greater than 0.10 includes 0.11, 0.12, 0.13, etc.). For example, whenever a numerical range with a lower limit, Rl, and an upper limit, Ru, is disclosed, any number falling within the range is specifically disclosed. In particular, the following numbers within the range are specifically disclosed: R=Rl+k*(Ru−Rl), wherein k is a variable ranging from 1 percent to 100 percent with a 1 percent increment, i.e., k is 1 percent, 2 percent, 3 percent, 4 percent, 5 percent, . . . , 50 percent, 51 percent, 52 percent, . . . , 95 percent, 96 percent, 97 percent, 98 percent, 99 percent, or 100 percent. Unless otherwise stated, the term “about” shall mean plus or minus 10 percent. Moreover, any numerical range defined by two R numbers as defined in the above is also specifically disclosed. Use of the term “optionally” with respect to any element of a claim means that the element is required, or alternatively, the element is not required, both alternatives being within the scope of the claim. Use of broader terms such as comprises, includes, and having should be understood to provide support for narrower terms such as consisting of, consisting essentially of, and comprised substantially of. Accordingly, the scope of protection is not limited by the description set out above but is defined by the claims that follow, that scope including all equivalents of the subject matter of the claims. Each and every claim is incorporated as further disclosure into the specification and the claims are embodiment(s) of the present invention.
Claims
1. A method of managing aircraft exhaust, comprising:
- receiving hot air from an exhaust aircraft engine at a hot air mass flow rate, the aircraft engine configured to drive a rotor of the aircraft;
- providing the hot air at the hot air mass flow rate to a mixer duct of an exhaust system attached to the exhaust duct;
- receiving cold air from outside the aircraft at a cold air mass flow rate through a first duct connecting the outside of the aircraft to a distributor disposed within the mixer duct, wherein a first fan is positioned within the first duct, and wherein the first fan is separate from the rotor;
- providing the cold air at the cold air mass flow rate to the distributor; and
- mixing the hot air and the cold air within the mixer duct at a variable hot air mass now rate to cold air mass flow rate ratio, wherein the hot air mass flow rate is varied using the aircraft engine and the cold air mass flow rate is varied by controlling a speed of the first fan positioned within the first duct, wherein the variable hot air mass flow rate to cold air mass flow rate ratio is selectively maintained independent of at least one of (1) a variation in the hot air mass flow rate and (2) a variation in a translational speed of the aircraft.
2. The method of claim 1, wherein the variation in the hot air mass flow rate is at least attributable to a variation in the engine speed.
3. The method of claim 1, wherein the variation in translational speed is at least partially attributable to a variation in an orientation of the engine exhaust system relative to a ground surface.
4. The method of claim 1, wherein the mixing comprises a variably controlled rate of film cooling.
5. The method of claim 1, wherein the mixing occurs in a space free of fan components.
6. The method of claim 1, wherein the mixing comprises preventing exhausted mixed air temperatures and surface temperatures of the exhaust system components that are visible from outside the aircraft from exceeding a selected temperature.
7. The method of claim 1, wherein at least some moving components of the first fan configured to selectively vary the cold air mass flow rate are not located in a path of unmixed hot air.
8. The method of claim 1, wherein the mixing comprises passing the hot air and the cold air through a substantially annular space defined by substantially round components.
9. The method of claim 1, wherein maintaining the variable hot air mass flow rate to cold air mass flow rate ratio reduces an infrared signature of the aircraft.
10. The method of claim 1, wherein the variable hot air mass flow rate to cold air mass flow rate ratio is equal to a value selected between 0.0 and 3.0, wherein the hot air has a hot air temperature, wherein a mixture comprising the hot air and the cold air has a mixture temperature, wherein the a ratio of the mixture temperature to the hot air temperature is equal to a value selected between 1.0 to about 0.338.
11. The method of claim 10, wherein at least one of the hot air and the cold air are supplied from multiple sources.
12. The method claim 1, further comprising maintaining the variable hot air mass flow rate to cold air mass flow rate ratio using one or more controllers.
13. The method claim 12, wherein maintaining the variable hot air mass flow rate to cold air mass flow rate ratio using the one or more controllers comprises selectively controlling the cold air mass flow rate based on factors comprising at least one of air temperature, surface temperature, pressure, or air flow.
14. The method of claim 1, further comprising:
- receiving cold air from outside the aircraft at a second cold air mass flow rate through a second duct connecting the outside of the aircraft to the mixer duct, wherein a second fan is positioned within the second duct, wherein the second fan is separate from the first fan and the rotor; and
- providing the cold air at the second cold air mass flow rate to the mixer duct; and
- wherein mixing the hot air and the cold air within the mixer duct at a variable hot air mass flow rate to cold air mass flow rate ratio comprises mixing in the mixer duct the provided hot air, the cold air received through the first duct, and the cold air received through the second duct, and wherein the second cold mass flow rate is varied by controlling a speed of the second fan positioned within the second duct.
15. An exhaust system for an aircraft engine, the exhaust system comprising an exhaust duct connecting an aircraft engine to a mixer duct attached to the exhaust duct, the exhaust duct configured to provide hot air from the aircraft engine at a hot air mass flow rate to the mixer duct, wherein the hot air mass flow rate is varied by the aircraft engine, and wherein the aircraft engine is configured to drive a rotor of the aircraft;
- a first duct connecting an outside of the aircraft to a distributor disposed within the mixer duct of the exhaust system, the first duct configured to receive cold air from outside the aircraft at a cold air mass flow rate and provide the cold air at the cold air mass flow rate to the distributor;
- a first fan positioned within the first duct and separate from the rotor; and
- a controller configured to control a speed of the first fan positioned within the first duct to provide the cold air through the first duct at the cold air mass flow rate to the distributor; and
- wherein the mixer duct is configured to mix the hot air from the exhaust duct and the cold air from the first duct at a variable hot air mass flow rate to cold air mass flow rate ratio, independent of at least one of (1) a variation in the hot-air mass flow rate and (2) a variation in a translational speed of the aircraft.
16. The exhaust system of claim 15, further comprising:
- a second duct connecting the outside of the aircraft to the mixer duct, the second duct being separate from the first duct, wherein the second duct is configured to receive cold air from outside the aircraft at a second cold air mass flow rate and provide the cold air at the second cold air mass flow rate to the mixer duct; and
- a second fan position within the second duct and separate from the rotor, wherein the controller is configured to control a speed of the second fan positioned within the second duct to provide the cold air through the second duct at the second cold air mass flow rate to the mixer duct.
17. The exhaust system of claim 15, wherein the distributor comprises a distributor internal space at least partially bounded by a distributor wall and at least one distributor perforation through the distributor wall connecting the distributor internal space with the mixer duct.
18. The exhaust system of claim 15, wherein the exhaust system is configured for use with a tiltrotor aircraft.
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Type: Grant
Filed: Sep 21, 2012
Date of Patent: May 10, 2016
Patent Publication Number: 20140084080
Assignee: Bell Helicopter Textron Inc. (Fort Worth, TX)
Inventors: Daniel B. Robertson (Southlake, TX), Dudley E. Smith (Arlington, TX), Robert M. Laramee (Fort Worth, TX)
Primary Examiner: Ehub Gartenberg
Assistant Examiner: Jared W Pike
Application Number: 13/624,465
International Classification: F02K 1/82 (20060101); F02C 7/24 (20060101); B64C 29/00 (20060101); F01D 25/30 (20060101); F02K 1/46 (20060101); F02K 1/34 (20060101);